CA2659448A1 - Wing panel structure - Google Patents
Wing panel structure Download PDFInfo
- Publication number
- CA2659448A1 CA2659448A1 CA002659448A CA2659448A CA2659448A1 CA 2659448 A1 CA2659448 A1 CA 2659448A1 CA 002659448 A CA002659448 A CA 002659448A CA 2659448 A CA2659448 A CA 2659448A CA 2659448 A1 CA2659448 A1 CA 2659448A1
- Authority
- CA
- Canada
- Prior art keywords
- wing panel
- outer layer
- inner layer
- panel structure
- core structure
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/26—Construction, shape, or attachment of separate skins, e.g. panels
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/40—Weight reduction
Abstract
A wing panel structure for an aerospace vehicle or the like may include a n outer layer of material having a predetermined thickness. A core structure may be placed on at least a portion of the outer layer of material. An inne r layer of material may be placed at least on the core structure. The inner layer of material may have a selected thickness less than the predetermined thickness of the outer layer of material.
Claims (39)
1. A wing panel structure for an aerospace vehicle, comprising:
an outer layer of material having a predetermined thickness;
a core structure placed on at least a portion of the outer layer of material;
and an inner layer of material formed at least on the core structure, wherein the inner layer of material has a selected thickness less than the predetermined thickness of the outer layer of material.
an outer layer of material having a predetermined thickness;
a core structure placed on at least a portion of the outer layer of material;
and an inner layer of material formed at least on the core structure, wherein the inner layer of material has a selected thickness less than the predetermined thickness of the outer layer of material.
2. The wing panel structure of claim 1, wherein the outer layer of material comprises a structure to predominantly support a wing load.
3. The wing panel structure of claim 1, wherein the outer layer of material comprises a multiplicity of plies of material.
4. The wing panel structure of claim 3, wherein the multiplicity of plies of material are cured and processed to a higher strength specification than the core structure and inner layer of material and are cured and processed before the core structure and inner layer of material are placed on the wing panel structure.
5. The wing panel structure of claim 3, wherein the multiplicity of plies of material comprise a multiplicity of epoxy unidirectional tape plies.
6. The wing panel structure of claim 5, wherein the multiplicity of epoxy unidirectional tape plies are cured and processed before placing the core structure and inner layer of material.
7. The wing panel structure of claim 3, wherein the plies of material are continuous for an extent of the wing panel.
8. The wing panel structure of claim 1, further comprising a layer of a non-destructive inspection (NDI) reflective material formed between the outer layer of material and the core structure.
9. The wing panel structure of claim 1, wherein the core structure comprises a honeycomb type structure.
10. The wing panel structure of claim 1, wherein the outer layer, the core structure and inner layer are assembled before curing.
11. The wing panel structure of claim 1, wherein the outer layer, the core structure and the inner layer are cured and processed to a higher strength specification.
12. The wing panel structure of claim 11, wherein the higher strength specification comprises curing at a temperature range between about 300 and about 400 degrees F and a pressure between about 80 and about 100 psi.
13. The wing panel structure of claim 1, wherein the inner layer of material comprises a plurality of plies of a fabric.
14. The wing panel structure of claim 1, further comprising:
a stiffener formed over at least the outer layer of material; and a support rib formed on the inner layer of material between the stiffener and an assembly including the core structure and the inner layer of material, wherein the inner layer of material extends under the support rib and overlaps a portion of a bottom flange of the stiffener.
a stiffener formed over at least the outer layer of material; and a support rib formed on the inner layer of material between the stiffener and an assembly including the core structure and the inner layer of material, wherein the inner layer of material extends under the support rib and overlaps a portion of a bottom flange of the stiffener.
15. The wing panel structure of claim 14, wherein the stiffener includes a group comprising an I section stiffener and a T section stiffener.
16. The wing panel structure of claim 14, wherein the stiffener is inboard of the assembly including the core structure and the inner layer of material.
17. A wing panel structure for an aerospace vehicle, comprising:
an outer layer of material having a predetermined thickness;
a core structure placed on a portion of the outer layer of material;
an inner layer of material formed at least on the core structure;
a stiffener placed on another portion of the outer layer; and a support rib placed on the outer layer of material between the stiffener and an assembly including the core structure and the inner layer of material.
an outer layer of material having a predetermined thickness;
a core structure placed on a portion of the outer layer of material;
an inner layer of material formed at least on the core structure;
a stiffener placed on another portion of the outer layer; and a support rib placed on the outer layer of material between the stiffener and an assembly including the core structure and the inner layer of material.
18. The wing panel structure of claim 17, wherein the outer layer of material comprises a structure to predominantly support a wing load.
19. The wing panel structure of claim 17, wherein the outer layer of material comprises a multiplicity of plies of material.
20. The wing panel structure of claim 19, wherein the multiplicity of plies of material are cured and processed to a higher strength specification than the core structure and inner layer of material and are cured and processed before the core structure and inner layer of material are disposed on the wing panel structure.
21. The wing panel structure of claim 17, further comprising a layer of a non-destructive inspection (NDI) reflective material disposed between the outer layer of material and the core structure.
22. The wing panel structure of claim 17, wherein the core structure comprises a honeycomb type structure.
23. The wing panel structure of claim 17, wherein the stiffener comprises a stringer of composite material.
24. The wing panel structure of claim 17, wherein the stiffener is inboard of the assembly including the core structure and the inner layer of material.
25. An aerospace vehicle, comprising:
a fuselage; and a wing extending from the fuselage, wherein the wing includes a plurality of wing panel structures, each wing panel structure including:
an outer layer of material having a predetermined thickness;
a core structure placed on at least a portion of the outer layer of material;
and an inner layer of material formed at least on the core structure, wherein the inner layer of material has a selected thickness less than the predetermined thickness of the outer layer of material.
a fuselage; and a wing extending from the fuselage, wherein the wing includes a plurality of wing panel structures, each wing panel structure including:
an outer layer of material having a predetermined thickness;
a core structure placed on at least a portion of the outer layer of material;
and an inner layer of material formed at least on the core structure, wherein the inner layer of material has a selected thickness less than the predetermined thickness of the outer layer of material.
26. The aerospace vehicle of claim 25, wherein the outer layer of material of each wing panel structure comprises a structure to predominantly support a wing load.
27. The aerospace vehicle of claim 25, wherein the outer layer of material of each wing panel structure comprises a multiplicity of plies of material and wherein the multiplicity of plies of material are cured and processed to a higher strength specification than the core structure and inner layer of material and are cured and processed before the core structure and inner layer of material are disposed on the wing panel structure.
28. The aerospace vehicle of claim 25, further comprising a layer of a non-destructive inspection (NDI) material disposed between the outer layer of material and the core structure of each wing panel structure.
29. The aerospace vehicle of claim 25, wherein the core structure of each wing panel structure comprises a honeycomb type structure.
30. A method of making a wing panel structure, comprising:
forming an outer layer of material having a predetermined thickness;
placing a core structure on at least a portion of the outer layer of material;
and forming an inner layer of material disposed at least on the core structure, wherein the inner layer of material has a selected thickness less than the predetermined thickness of the outer layer of material.
forming an outer layer of material having a predetermined thickness;
placing a core structure on at least a portion of the outer layer of material;
and forming an inner layer of material disposed at least on the core structure, wherein the inner layer of material has a selected thickness less than the predetermined thickness of the outer layer of material.
31. The method of claim 30, wherein forming the outer layer of material comprises forming a structure to predominantly support a wing load.
32. The method of claim 30, wherein forming the outer layer of material comprises:
depositing a multiplicity of plies of material;
curing and processing the multiplicity of plies of material to a higher strength specification than the core structure and inner layer of material.
depositing a multiplicity of plies of material;
curing and processing the multiplicity of plies of material to a higher strength specification than the core structure and inner layer of material.
33. The method of claim 32, wherein the multiplicity of plies of material of the outer layer of material are cured and processed before the core structure and inner layer of material are disposed on the wing panel structure.
34. The method of claim 30, further comprising forming a layer of NDI
reflective material between the outer layer of material and the core structure.
reflective material between the outer layer of material and the core structure.
35. The method of claim 30, wherein placing the core structure comprises placing a honeycomb type structure.
36. The method of claim 30, wherein forming the inner layer of material comprises laying a plurality of plies of a fabric.
37. The method of claim 30, further comprising:
placing a stiffener over at least the outer layer of material; and placing a support rib on the inner layer of material between the stiffener and an assembly including the core structure and the inner layer of material, wherein the inner layer of material extends under the support rib and overlaps a portion of a bottom flange of the stiffener.
placing a stiffener over at least the outer layer of material; and placing a support rib on the inner layer of material between the stiffener and an assembly including the core structure and the inner layer of material, wherein the inner layer of material extends under the support rib and overlaps a portion of a bottom flange of the stiffener.
38. The method of claim 30, further comprising curing the wing panel structure after forming the inner layer of material.
39. The method of claim 38, wherein curing the wing panel structure comprises applying a temperature between a range between about 300 and about 400 degrees F and a pressure between about 80 and about 100 psi.
Applications Claiming Priority (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US11/553,017 US7628358B2 (en) | 2006-10-26 | 2006-10-26 | Wing panel structure |
| US11/553,017 | 2006-10-26 | ||
| PCT/US2007/016377 WO2008105806A2 (en) | 2006-10-26 | 2007-07-18 | Wing panel structure |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CA2659448A1 true CA2659448A1 (en) | 2008-09-04 |
| CA2659448C CA2659448C (en) | 2012-06-19 |
Family
ID=39328954
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA2659448A Active CA2659448C (en) | 2006-10-26 | 2007-07-18 | Wing panel structure |
Country Status (8)
| Country | Link |
|---|---|
| US (1) | US7628358B2 (en) |
| EP (1) | EP2076431B1 (en) |
| JP (1) | JP5319538B2 (en) |
| CN (1) | CN101557979B (en) |
| CA (1) | CA2659448C (en) |
| ES (1) | ES2770642T3 (en) |
| PT (1) | PT2076431T (en) |
| WO (1) | WO2008105806A2 (en) |
Families Citing this family (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20080277531A1 (en) | 2007-05-11 | 2008-11-13 | The Boeing Company | Hybrid Composite Panel Systems and Methods |
| IT1392320B1 (en) * | 2008-12-09 | 2012-02-24 | Alenia Aeronautica Spa | ATTACK EDGE FOR WINGS AND AIRCRAFT MAKES |
| DE102009013585B4 (en) * | 2009-03-17 | 2012-01-26 | Airbus Operations Gmbh | Fuselage cell structure for a hybrid aircraft |
| US8167245B1 (en) * | 2009-11-03 | 2012-05-01 | The Boeing Company | Fuel barrier |
| JP5535957B2 (en) * | 2011-02-21 | 2014-07-02 | 三菱航空機株式会社 | Formation method of wing panel |
| US9943937B2 (en) | 2012-09-28 | 2018-04-17 | The Boeing Company | System and method for manufacturing a wing panel |
| EP2989003A4 (en) * | 2013-04-25 | 2016-12-07 | Saab Ab | Stiffening element run-out |
| US10801836B2 (en) | 2017-06-13 | 2020-10-13 | The Boeing Company | Composite parts that facilitate ultrasonic imaging of layer boundaries |
Family Cites Families (14)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2382358A (en) * | 1944-02-03 | 1945-08-14 | Budd Edward G Mfg Co | Stressed skin airfoil joint |
| US3058704A (en) * | 1958-01-16 | 1962-10-16 | Johnson & Johnson | Laminated adhesive sheeting for aircraft |
| US4344995A (en) * | 1980-09-15 | 1982-08-17 | The Boeing Company | Hybrid composite structures |
| US4599255A (en) * | 1981-12-28 | 1986-07-08 | The Boeing Company | Composite structures having conductive surfaces |
| US4542056A (en) * | 1983-08-26 | 1985-09-17 | The Boeing Company | Composite structure having conductive surfaces |
| WO1985001489A1 (en) * | 1983-09-29 | 1985-04-11 | The Boeing Company | High strength to weight horizontal and vertical aircraft stabilizer |
| DE19529476C2 (en) * | 1995-08-11 | 2000-08-10 | Deutsch Zentr Luft & Raumfahrt | Wing with shear-resistant wing shells made of fiber composite materials for aircraft |
| US5866272A (en) * | 1996-01-11 | 1999-02-02 | The Boeing Company | Titanium-polymer hybrid laminates |
| JP2000043796A (en) * | 1998-07-30 | 2000-02-15 | Japan Aircraft Development Corp | Wing-shaped structure of composite material and molding method thereof |
| DE19845863B4 (en) * | 1998-10-05 | 2005-05-19 | Deutsches Zentrum für Luft- und Raumfahrt e.V. | Structural element with large unidirectional stiffness |
| US6976829B2 (en) * | 2003-07-16 | 2005-12-20 | Sikorsky Aircraft Corporation | Rotor blade tip section |
| US7115323B2 (en) * | 2003-08-28 | 2006-10-03 | The Boeing Company | Titanium foil ply replacement in layup of composite skin |
| US7052573B2 (en) * | 2003-11-21 | 2006-05-30 | The Boeing Company | Method to eliminate undulations in a composite panel |
| US7325771B2 (en) * | 2004-09-23 | 2008-02-05 | The Boeing Company | Splice joints for composite aircraft fuselages and other structures |
-
2006
- 2006-10-26 US US11/553,017 patent/US7628358B2/en active Active
-
2007
- 2007-07-18 ES ES07873811T patent/ES2770642T3/en active Active
- 2007-07-18 EP EP07873811.9A patent/EP2076431B1/en active Active
- 2007-07-18 CA CA2659448A patent/CA2659448C/en active Active
- 2007-07-18 JP JP2009534567A patent/JP5319538B2/en active Active
- 2007-07-18 WO PCT/US2007/016377 patent/WO2008105806A2/en active Application Filing
- 2007-07-18 CN CN2007800364237A patent/CN101557979B/en active Active
- 2007-07-18 PT PT78738119T patent/PT2076431T/en unknown
Also Published As
| Publication number | Publication date |
|---|---|
| CN101557979B (en) | 2012-06-13 |
| WO2008105806A3 (en) | 2009-06-11 |
| JP5319538B2 (en) | 2013-10-16 |
| US7628358B2 (en) | 2009-12-08 |
| WO2008105806A2 (en) | 2008-09-04 |
| JP2010507530A (en) | 2010-03-11 |
| CN101557979A (en) | 2009-10-14 |
| PT2076431T (en) | 2020-01-09 |
| US20080099613A1 (en) | 2008-05-01 |
| ES2770642T3 (en) | 2020-07-02 |
| EP2076431B1 (en) | 2019-12-04 |
| CA2659448C (en) | 2012-06-19 |
| EP2076431A2 (en) | 2009-07-08 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| EEER | Examination request |